Characteristics of composite materials have meant that composite components are employed in an increasing range of applications from aerospace to automotive parts.
In the aerospace industry, for example, composite materials have been used for a number of years owing to their strength to weight ratio. The term “composite materials” (known also as “composites”) is used to describe materials comprising for example glass fibre or carbon fibres and an epoxy resin (or similar). These are also known as glass reinforced plastic or carbon fibre reinforced composites. The carbon fibre reinforced composite material offers improved properties such as lower weight, improved fatigue/damage resistance, corrosion resistance and negligible thermal expansion.
The use of these materials has increased throughout the aerospace industry predominantly because of the fuel savings which can be achieved over the life of an aircraft by reducing the overall sum weight of the components making up the aircraft. Aerodynamic as well as structural components are formed of composite materials and particularly carbon fibre materials.
A composite component may be laid-up using a cloth, tape or the like pre-impregnated with resin to form a stack corresponding to the desired shape of the part to be formed. The stack is then cured either at ambient temperature and pressure or at elevated temperature and pressure in an autoclave to create a hardened component.
A gas turbine engine such as a turbofan may be provided with a containment case for preventing a broken blade of the engine from exiting the engine and damaging the rest of the aircraft. For example, a containment case may be provided around the fan at the front of the turbofan engine. The containment case may be made of composite material such as carbon fibre reinforced composite material and/or Kevlar reinforced composite material. The containment case is in the shape of a generally cylindrical barrel or housing. The containment case needs to be attached to adjacent structural components of the engine and it is therefore desirable for the containment case to include a flange at one or both of the ends of the barrel or housing.
FIG. 1 is a diagrammatic side view of a typical known turbofan engine 1 having a fan case 11 defining a fan duct 12 which contains a rotating disc of fan blades 13. The fan blades 13 rotate around a central longitudinal axis 14 of the engine 1.
The fan case 11 is annular and is centred on the longitudinal axis 14. The fan case 11 is shown partly cut away in FIG. 1 in order to diagrammatically illustrate the fact that the fan case 11 includes an annular containment case 2 positioned around the periphery of the disc of fan blades 13 in order to contain any broken fan blade 13. The containment case 2 comprises a generally-cylindrical barrel or housing 3 at the front end of which is an outwardly-extending annular flange 41 and at the rear end of which is an outwardly-extending annular flange 42.
The containment case 2 is centred on the longitudinal axis 14 of the engine 1 and is held in position by being fastened to other components of the fan case 11 such as an annular front leading edge 51 and an annular rear edge 52. The flanges 41 and 42 may be provided with holes for fasteners which are used to attach the containment case 2 to the structure of the leading and rear edges 51, 52.
FIG. 2 is a diagrammatic perspective view of a containment case 2 generally similar to the one shown in FIG. 1 except that in FIG. 2 the generally-cylindrical housing 3 is slightly tapered in the direction of the longitudinal axis 14. The actual internal contour or profile of the housing 3 may be optimised to suit the requirements of a particular engine 1.
It is convenient to use a machine, such as an automated tape laying (ATL) machine, to lay-up the plies of composite material of the housing of the containment case on a mould or mandrel. It has proved difficult to use a machine to lay-up the plies of the composite material of the flange and to integrate the composite material of the flange with the composite material of the housing, before the housing and the flange are cured. It has proved necessary to manually lay-up the plies of the flange, ply by ply, against an outwardly-projecting annular wall of the mould which extends outward from the main cylindrical mould surface on which the plies of the housing have been machine-laid up. The plies of the flange are hand laid and must be intermeshed with the machine-laid plies of the housing. This tends to produce a flange of inconsistent quality and, in order to compensate for this, a flange which is heavier than it needs to be because it is using an excess of composite material.
At GKN Aerospace, we have recently been experimenting with a configuration during the laying-up of the composite material which facilitates machine laying of the flange in addition to machine laying of the curved main surface of a composite structure. In this way, all of the laying-up of the composite material may be automated, and it is no longer necessary to use manual or hand laying. Automating the laying-up of the flange produces an improvement in the quality of the flange and the composite structure.
Our experimental tool for forming (bending) a second part 43 to form a flange 4 (corresponding, for example, to part of the front flange 41) is shown in FIGS. 3 to 6. The experimental tool incorporates a first part 61 of a cylindrical mould or mandrel 6 and a circumferential line of movable blocks 62. An experimental part 21 of the containment case 2 is laid-up as a preform on a mould surface 611 of the first mould part 61 and on mould surfaces 621 of the movable blocks 62. The preform comprises the second part 43 and a first or main part 31. The second part 43 is laid-up on the mould surfaces 621 and the first part 31 is laid-up on the mould surface 611.
Pre-impregnated uni-directional tape making up the composite material of the experimental part 21 is laid-up obliquely (e.g. at plus 60° and at minus 60° relative to the circumferential direction) in both the first part 31 and the second part 43. In the context of the experimental part 21, the circumferential direction corresponds to the boundary line 211 between the parts 31, 43. Circumferential tape (0° tape) is laid-up in the first part 31 but not in the second part 43. Perpendicular tape (90° tape) is laid up in the second part 43 and extends a short distance into the first part 31.
For example, the tape is 0.25 mm thick and has a typical width of 75 mm to 150 mm. Such tape is suited to being laid-up by the head of an ATL machine. The tape is laid-up to form a stack of plies, and the number of ply layers may be 10 or more, preferably 20 or more, or preferably 30 or more.
The laid-up tape of the experimental part 21 is cut back so that the circumferential free edge 431 of the second part 43 does not project beyond movable blocks 62 of the mould 6.
A female forming tool 63 (see FIG. 4) is then clamped down onto the part of the first part 31 immediately adjacent the second part 43.
The mould 6 is then placed in an oven and heated to a first temperature, e.g. 80° C., at which the resin of the pre-preg tape becomes fluid enough (has a low enough viscosity) to facilitate the forming operation which is about to occur. At this point, the blocks 62 are advanced in a radially outward direction from their retracted or flush position to the advanced position shown in FIGS. 5 and 6 by actuating actuators 64 such as pneumatic or hydraulic pistons. This forms or flexes upwards the second part 43 to form the flange 4 projecting outwards relative to the first part 31. The blocks 62 are advanced by a distance which is at least the width 432 of the second part 43 so that the flange 4 that is formed is an upright wall relative to the first part 31.
The temperature in the oven is then raised to a second, higher temperature in order to continue and complete the curing of the composite tape material. For example, the second temperature may be 135° C. After the curing cycle or process has been completed, the blocks 62 may be retracted to the position shown in FIGS. 3 and 4. The female tool 63 may be removed, and the experimental part 21 removed from the mould 6.
FIGS. 7 and 8 are diagrammatic cross-sectional views through the flange wall 4 (the second part 43) and the adjacent part of the first part 31 after the forming operation has been performed, and FIG. 7 shows the outcome of a correct forming operation and FIG. 8 shows the outcome of an incorrect forming operation.
If the forming operation is correctly performed as shown in FIG. 7, the plies 44a-44g slide over one another in the second part 43 and the thickness of the second part 43 is not distorted. However, a so-called “bookend” effect 433 is produced at the free edge 431 of the second part 43, whereby the free edge 431 is slanted.
The flange 4 may be trimmed to a desired height along a cut line 434 and this will remove the unwanted bookend effect 433.
In FIGS. 7 and 8, only 7 plies 44a-44g are shown for reasons of clarity, but in practice it is usually the case that a larger number of ply layers will be present.
During the forming operation, the bending upwards of the second part 43 will tend to place the plies at the inside of the bend (such as plies 44a and 44b) in compression and will tend to place the plies at the outside of the bend (such as plies 44f and 44g) in tension.
If the forming operation is performed too quickly, such that the plies are unable to move relative to one another to the necessary extent, a flange 4 as shown in FIG. 8 may be the end result. The forming (bending) has been performed too quickly, and the compression of the plies at the inside of the bend (such as plies 44a and 44b) has resulted in them undergoing wrinkling or buckling as shown at area 435. Even after the flange wall 4 has been trimmed to height along the cut line 434, some of the unwanted distortion 435 will still remain.
It would be desirable to improve the manufacturing method so as to reduce or eliminate this unwanted distortion.